Mirror landing system



Oct. 1o, 1961 5 Sheets-Sheet 1 Filed Sept.v 10, 1959 HTTO/ZA/e-YS Oct10, 1961 J. A. LUNDIN ETAL 3,003,451

' MIRROR LANDING SYSTEM Filed Sept. 10, 1959 5 Sheets-Sheet 2 ATToRA/EYSA JOI-IN A, LUND/A/ I *l Q GEORGE P. MSELL/ R 5' Sheets-Sheet 3 Oct. 10,41.961 J. A. LuNDlN ETAL 'R MIRROR LANDING SYSTEM med sept. 1o, 1959Oct'. l0, 1961 J. AQ I UNDIN ETAL MIRROR LANDING SYSTEM 5 Sheets-Sheet 5Filed Sept. 10, 1959 JOHN A. UND/N GEOQE RMASELL/ L/ENQYQZUL-'PNDoQ/ffzINVENTOR. MQW z 6m BY ffy/77%] eyfd QS. mmu E @mi m G wil! wwwm QmvfwAll! Sow 3,093,451 MIRROR LANDING SYSTEM .lohn A. Lundin, Madison,George P. Maselli, Dumont,

and Henry D. Zuerndorfer, Pompton Plains, NJ., as-

signors to @enteral Precision Inc., a corporation of Delaware FiledSept. 1li, 1959, Ser. No. 839,241 5 Claims. (Cl. IIL-43.5)

lThis invention relates to mirror landing systems for aircraft and, moreparticularly, to a novel mirror landing system including stabilization.means.

Mirror landing systems for aircraft, which are well known in the art,may be used both. on land and sea to aid an aircraft in landing, along apredetermined glide path. However, they `are particularly suited for usein landing an aircraft on the deck of a ship, such as an aircraftcarrier.

Briefly, present day mirror landing systems comprise a plane mirrorlocated on the deck of a ship facing a landing 'aircraftA On either sideof the mirror are datum lights oriented midway between the top' andbottoni of thev mirror which also face the aircraft. A source of lightlocated on the deck oi' the ship is directed toward the mirror andreflected therefrom. Assuming that the ship is not subject to pitch,roll or heave, the angle of incidence of the source light is such. thatthe virtual image thereof, as seen. from the aircraft,r will appear atthe center of the mirror, midway between the datum lights and in linetherewith, only if the aircraft is on a predetermined glide path. lt-Ethe aircraft. should be above the predetermined glide path, the virtualimage will appear above the datum lights, and if the aircraft should` bebelow the predetermined glide path, the virtual image will he below thedatum lights.

However,v the ship is, in. fact, very much subject, to pitch, roll andheave. Therefore the virtual image, as seen from the aircraft, appearsto dance about a point above, below or in line with'the` datum lights,as the case may be. This seriously reduces the eifectiveness of presentday mirror landing systems.

One proposed arrangement for overcoming this problem is to. maintain themirror relatively' fixedv in space. ln order to accomplish this it isnecessary to mount the mirror on an elevator which is moved in responseto pitch,v roll and heave. This solution is not too practical, since itinvolves structural changes in an aircraft. carrier. Furthermore, due tothe large inertia of such an elevator, it is doubtful that it wouldeffectively accomplish its purpose.

The present invention contemplates a relatively simple arrangement forautomatically tiltingl the mirror in accordance with pitch, roll"` andheave in suchV a manner that the virtual image, as seen from theaircraft, remains centered between the datum lights, if the aircraft ison the predeterminedv glide path, despite Vsuch pitch roll and heave.

lt is therefore. an. objject of the present' invention to provi'd'e animproved mirror landing system.

lt is a further object of this invention to. provide a mirror llandingsystem incorporating, stabilization means.

It is a still further objectv of this invention to, provide automaticmeans for'til-ting the mirrorY of. a mirror landing system.

It is one feature of the present invention to' provide an infraredtracking system for determining the elevation mirror tilti angle in.response to an input from. the track.-

k.ted States Patent 0 icc ing system manifesting elevation angle andother inputs from the ships gyros manifesting pitch and roll.v

These and other objects, features and advantages of the presentinvention will becomev apparent from the' fol'- rlowing detaileddescription taken together with the accompanying drawing, in which:

FlG. l is a diagram of an aircraft carrier and landing aircraftutilizing the mirror landing system of' this in'- vention,

FIGS. 2A and 2B illustrate the geometric relationships which exist in'this mirror landing system,

FlG. 3` is a block diagram of the major components of.' the mirrorlanding system,

FIG. 4- is.' a block diagram of the Atilt angle computer of FIG'. 3,

FIG. 5" is a block diagraml of the infra-red tracker of FIG. 3, and

FG'. 6 is a biock Iand schematic diagram of the: height change computerof FIG; 3.

Referring now to FIG. l, there is shown an aircraft carrier it). Mountedon the deck of aircraft carrier 10@ is mirror 102 with datum lights 1.64and light source 106. The light from source ldd is reected from mirror102. Light source 166 is oriented at a fixed angle with respect to thedeck of aircraft carrier lili); However, the angle ci incidence, as wellas. the angle of reflectiom is varied by the tilt angle ot. mirrorlttZ.V

Infra-red sensing device 103, located just. under the deck at the how ofaircraft carrier 100,` provides information as to the value ot angle abetween track line and line llZ parallel to the deck of aircraft carrieritiil. Angle along with pitch and roll information obtained romY thegyros of aircraft carrier lllll are used to tilt mirror lli?. to suchvan angle that the line of sight lle from landing aircraft llo tothevirtual image of light source lila passes through. the center of mirror102 only when landing aircraft` 1116 is located at the intersection of`tracking line l1@ and predetermined glide path 118.

In this case, as shown in FIG. l', virtual image 12d of .light sourcei016 appears in line with datum lights 104 and midway therebetween.

Should the landing aircraft be located on tracking line 110 above theintersection. thereof with predetermined glide path 118, as shown at llaof FIG. l, the line of sight lido from `aircraft llda to the virtualimage of light source 166 will pass through mirror 102 above the centerthereof. Therefore, as shown in FIG. l, virtual image l2tla appearsabove datum lights 1M.

In a similar manner, should the landing aircraft be located'. ontracking line lill below the intersection thereof'with predeterminedglide path 11.18', as shown at H613 of' FIG. l, the line of sight fromaircraft ldb to the virtual image of light source luc. will pass throughmirror 102 below the center thereof. Therefore, as shown in FIG.Y l',virtual image 126i? yappears below datum lights 104'.

FIGS. 2A'. and 2B show the geometric relationships which exist'indeterminingv the mirror tilt angle p..

nFGS. 2A and 2B, M` represents the actual. position of thecenter ofmirror lli'Zi, M represents the position the center ofthe mirror wouldhave if: there were' Zero pitch roll and heave, a represents the fixeddistance hetween Mv and M', A It' represents the vertical'. compmient ofa, line M4" represents the line of sight between landing aircraft. lida-ndM, line lid represents the predetermined glide. path, line Elli)represents the projection of line 114 on. the horizontal plane throughM', e represents the angle between lines lili, andZd, and' reprcsentsthe pitch. angle..

ln FIG. 2A, L represents the position ofA light source 106, line4 202.represents a line parallel to the deck 3 through L, line 204 representsthe incident beam of light directed toward M from L, represents thelight source angle between lines 202 and 204, line 206 represents a lineparallel to lthe deck through M, line 208 represents the perpendicularto the deck through M, line 210 represents the vertical through M, 4andV represents the position of the virtual image of light source 106 asseen from landing aircraft 116.

In FIG. 2B, P represents the position of landing aircraft 116, Irepresents the fixed distance between I and M, line 212 represents aline parallel to the deck through l, line 110 represents the trackingline between P and I, a represents the elevation angle between trackingline 110 and line 212, 'y represents the fixed angle between b and line212, B represents the angle between tracking line 110 and line of sight114, A represents the angle between line of sight 114 and predeterminedglide path 118, e represents the angle between a and Ah, Ax representsthe yhorizontal component of a, line 214 represents the horizontal datumline through M for Ah equal to zero, line 216 represents the deck ofaircraft carrier 100 for Ah equal to zero, and qa represents the glideangle between predetermined glide path 118 and line 216.

From consideration of FIG. 2A it can be shown that qs arc tan Therefore,

(3) acos (qst-aga cos (--e)=a cos e=Ah (4) Sin qs'qs SubstitutingEquations 3 and 4 in Equation 2,

Ah, a, b, Iy and gb' are all known or measurable quantities. Therefore,qt may be ascertained with insignificant error by solving Equation 5.After solving Equation 5, it is a simple matter to solve Equation 1 toobtain the tilt `angle n.

Referring now to FIG. 3, there is shown a block diagram of a preferredembodiment of the invention. Tilt angle computer 302, which is shown indetail to FIG. 4, includes handwheel 304 for setting shaft input 306 toa position qt', corresponding to the value of glide angle, and,handwheel 308 for setting shaft input 310 to a position -6 correspondingto the value of the negative of the light source angle.

Infra-red tracker 312, shown in detail in FIG. 5, provides a rstelectrical output proportional to elevation angle a which is applied asan input to tilt angle computer over conductor 314, and a secondelectrical output proportional to b sin (a4-7) which is applied as aninput to tilt angle computer 302 over conductor 316.

Height change computer 318, shown in detail in FIG. 6, which has a firstelectrical input proportionate to roll angle )t and a second electricalinput proportional to vpitch angle applied thereto from the ships gyrosover conductors 320 and 322, respectively, provides an elec- 4 tricaloutput proportional to Ah which is applied as an input to tilt anglecomputer 302, over conductor 324.

An electrical input proportional to pitch angle from the ships gyros isalso directly applied as an input to tilt angle computer 302 overconductor 326.

Tilt angle computer 302, in response to the inputs applied thereto,solves Equations 5 and l, discussed above, and produces an electricaloutput proportional to mirror tilt angle p, which is applied as an inputto tilt angle servo system 328 over, conductor 330.

In response to this input being applied thereto, tilt angle servo system328 rotates output shaft 332 thereof, which is coupled to mirror anddatum lights 102, to thereby tilt mirror and datum lights 102 throughthe angle p..

Referring now to FIG. 4, which is a detailed showing of tilt anglecomputer 302, pitch follow-up servo 402 is responsive to pitchinformation. More specifically, pitch angle infomation is transmitted at2 and 36 times pitch angle from the ships gyros by respective synchrotransmitters associated therewith. The 2 times pitch angle informationis applied as an input to control transformer 404 and the 36 times pitchangle information is applied as an input to control transformer 406.

The respective outputs of control transformers 404 and 406 are connectedas respective inputs to crossover network 408. The output of crossovernetwork 408 is applied as a iirst input to servo amplier 410. The outputof servo amplifier 410 energizes servo motor 412 of motor-generator set414.

Servo motor 412 drives generator 416 and rotates shaft 418. The outputof generator 416 is fed back as a second input to servo amplifier 410.

Shaft 41S is coupled through suitable reduction gearing 420 and shaft422 to the rotor of control transformer 406. Shaft 422 is also coupledthrough 18/1 reduction gearing 424 and shaft 426 to the rotor of controltransformer 404. Shaft 426 is also coupled through 2/1 reduction gearing428 to shaft 430.

In response to the operation of servo motor 412, the rotors of controltransformers 404 and 406 are driven to null positions, at which pointservo motor 412 stops. At null, the angular position of shaft 430relative to a reference position is equal to pitch angle Elevation anglea information is transmitted by a synchro transmitter associated withinfra-red tracker 312 and is applied as an input to differential synchrotransmitter 432.

Since the rotor of differential synchro transmitter 432 is oriented atpitch angle 6 relative to the stator thereof, the output therefrom,which is applied as an input to control transformer 434 of tracker servo436 manifests in electrical form the value of (or-).

In the manner described in connection with pitch follow-up servo 402,the servo loop of tracker servo 436 comprising servo amplifier 438,motor-generator set 440 and reduction gear 442, serves to drive controltransformer 434 to null. At null, the angular position of shaft 444relative to a reference position thereof is equal t0 (aL-).

Shaft 444 is used to rotate the rotor of resolver 446 to an angularposition of (at-) relative to the stator thereof.

A signal manifesting Ah, which is received from height change computer318, is applied as an input to resolver 446. Therefore resolver 446 willproduce a first output lequal to Ah sin (ar-), which is applied to afirst input apparsi fore, the, output from linear potentiometer 452,which is applied as a second input to summing network 448, is equal to bsin (n+1/)df and the output from cosine potentiometer 454, which isapplied as a second input to summing network 450, is equal to b sin(ot-l-fy) cos gb.

'The outputs. of summing networks 448 and 450 which are respectivelyequal to Ah sin (a-H-b sin (e4-10p" and Ah cosl (.--)+b sin (a-l-fy) costp are applied as respective first and second inputs to resolver 456 oftangent servo 458; The output from resolver 456 is applied through aservo loop comprising servo amplifier 460, motor-generator set 462,reduction gear` 464 and shaft 466, to drive the rotor of resolver 456,which is coupled to shaft 466', to a null position. At null the angularposition of shaft 466 relative to a reference position is equal to Allsin (nz-6) -l-b sin (arl-7) qb' Alk cos (pr-) -l-bsin (a+-y) cos qbThis, it will be seen from Equation 5, is equal to fp.

Shaft 466 is connected as a first input to differential 468 and shaft430, having an angular position equal to is connected as a second inputto differential 46S. Gutput shaft 470 of differential 468 therefore hasan angular position relative to a reference position equal to (qb-l-).Shaft 470 is connected as a rst input to differential 472. 9, thenegative ofl light source angle, is applied as a second input to`differential 472 through handwheel 308 and shaft 310.

Therefore, the angular position of output shaft 474 relative to areference position is equal to (-l,6). However, from Equation 1,

Thus, the angular position of shaft 474 is equal to twice the mirrortilt angle p..

Shaft 474 is connected directly to the rotor of synchro transmitter `476and through I/'lS step-up gear 478 to synchro transmitter 480.`

The outputs from synchro transmitters 476 and 480, which manifestrespectively 2 and 36 times the mirror tilt angle n, are applied asinputs to tilt angle servo system 328 of FIG. 3. In response to thisinput information, tilt angle servo system 328 is effective in tiltingmirror and datum lights 102 of FIG. 3 through the mirror tilt angle p..

Referring now to FIG.. 5,. which is a Idetailed showing of infra-red.ltracker 312 of FIG. 3, infra-red sensing device 108. of FIG. 'Zcomprises a folded reiiector optical system composed of reflectors 502and 504 for focusing incident infra-red radiation on infra-red sensor506 through reticle 503i Reticle 50S? is rotated at a given xedfrequency by motor 510. Located near the reticle is an appropriatelyplaced azimuth pickofr' 512 and an appropriately placed elevationpickoff 514. "E

In response to the rotation of reticle 508 a signal is induced inazimuth pickoff 512 which has a frequency equal to that of reticle 50Sand a phase determined by the position of azimuth pickoff 512. In asimilar manner, a signal is induced in elevation pickolf 514 which has afrequency equal to that of reticle 508 and a phase determined by theposition of elevation pickoff 514.

The signal from azimuth pickofi 512 is applied as a first input toazimuth phase discriminator 516 and the signal from elevation pickolf514 is applied as a first input to elevation phase discriminator 518.

The output from infra-red sensor 506 is a signal having a frequencyequal to that of reticle 508 and a phase determined by the orientationof infra-red sensing device 108 with respect to the direction of alanding aircraft. The signal from infra-red sensor 506 is amplified byamplifier 520 and applied as a second input to both azimuth andelevation phase discriminators 516 and 518.

are tan Azimuth phase discriminator 516 produces an output which isVproportional to the phase difference between the phase of the' secondinput thereto frominfra-red sensor 5061 and -thephase of the first inputtheretoy from azimuth pickoff 512.

The output from azimuth phase discriminator 516- is applied to anazimuth servo loop comprising servo antplifier 522, motor generator set524, reduction gear 526 and shaft 52S. Shaft 52S is connected toinfra-red sensing device 100, and serves to rotate infra-red sensingdevice 108 in azimuth about its center to a null point at which there iszero phase difference between the first and second inputs to azimuthphase discriminator 5116.

In a similar manner', elevation phase discriminator 518 and the servoloop emanating therefrom comprising servo amplifier 530, motor-generatorset 532, reduction gear 534, and shaft 536 serves toy rotate infra-redsensing devi-ce 108 in elevation about its center to a null point atwhich, there zero phase difference between the first and second inputsto elevation phase discriminator 5181.

When both azimuth and elevation phase discriminators S16 and 51S` are attheir null points, infra-red sensing device 10Sv points directly' at thelanding aircraft.

Coupled to elevation shaft 536- is synchro transmitter 5318l fortransmitting a signal manifesting elevation angle a tol tilt anglecomputer 302 of FIG. 3.

Elevation shaft 536v is alsoy coupled to the rotor of resolver 538. Therotor of resolver 538 is offset' relative tov the stator thereof by thefixed angle fy. An input Voltage equal to I:- (not shown) is connectedas an input resolver 538. Therefore, resolver 53S produces an outputequal tob sin (a4-v), which is applied to tilt angle' computer 302 ofFIG. 3.f

Referring nowto FIG. 6, which is a detailed showing of height changecomputer 318 of FIG. 3, vertical accelerometer 692 is mounted in innerpitch gimbal' 604 which is supported by roll gimbal 606 for verticalstabilization. Pitch gimbal 604 is controlled by pitch angle informationtransmitted thereto from a synchro transmitter associated with the shipsgyros. Roll gimbal 606- is controlled by roll angle 7i informationtransmitted thereto from a synchro transmitter also associated with theships gyros.

Associated with vertical' accelerorneter 602 is a torquer 608 and apickoif 610. Pickoif 610 has an excitation signal applied thereto from4000 c.p.s'. source 612. In response to vertical acceleration, pickoff610 produces a 4,000 c.p.s. signal having an amplitude proportional tothe vertical acceleration which is applied as an input to amplifierAmpliher 614-; which has a second 4000 c-.p.s` signal applied theretoIas a reference signal directly from` 4000- c.p'.s-., source 612,.produces: a D.C. voltage output having at magnitude proportional. to`the vertical acceleration.

The output of amplifier 614 is applied in series with torquer 608 as aninput to first integrator 616, which comprises resistances 618 and 620,servo amplifier 622, servo motor 624, reduction gear 626, shaft 628,voltage divider 630 operated by shaft 628, and capacitance 632.

The output of first integrator 616, which is proportional to verticalvelocity, is applied as an input to second integrator 634, which iscomposed of resistance 636, servo amplifier 638, servo motor 640,reduction gear 642, shaft 644, voltage divider 646 which is operated byshaft 644, and capacitance 648.

Shaft 644 also operates the wiper of A.C. voltage divider 650. Thevoltage at the wiper of voltage divider 650 is equal to the heightchange Ah, `and is applied to -tilt angle computer 302 of FIG. 3.

In order to maintain the accuracy of first and second integrators 616an-d 634 over long periods of time both proportional and integralfeedback loops are incorporated in height change computer 318.

More specifically, the A C. output on the wiper of voltage divider' 650,equal to Ah, is applied 'as a lirst input to differential network 652.An A.C. signal having an amplitude representing a height referencehorizontal datum line is applied as a second input to differen-tialnetwork 652. The output from differential network 652, equal to thedifference between the reference height and Ah, is applied to aproportional feedback network comprising demodulator 654, and is furtherapplied to an integral feedback network comprising servo amplifier 656,motor-generator set 658, reduction gear 660, shaft 662, and voltagedivider 664, which is operated by shaft 662.

The output from demodulator 654 is applied through series resistance 666to load resistance 668, and the output from voltage divider 664 isapplied through series :resistance 670 to load resistance 663. Theoutput across load resistance 668 is applied as a feedback input tofirst integrator 616 through resistance 672, and is applied as afeedback input to second integrator 634 through re- `sistance 674. Thusi-t will be seen that Ah is equal to the double integral of the verticalacceleration measured by the vertically stabilized verticalaccelerometer. In order to ensure that Ah represents the true change inheight experienced by mirror and `datum lights 102 due to pitch, rolland heave, vertical accelerometer is mounted in close proximity tomirror and datum lights 102.

Although only a preferred embodiment of the invention has been describedherein, it will be apparent there are modifications and changes whichfall within the skill of the art. For instance, there are trackingmethods other than infra-red, such as radio and radar, for example. Alsoother circuit arrangements than that shown may be used to'solveEquations l and 5. It is, therefore, not intended that the invention berestricted to the preferred embodiment disclosed herein, but that it belimited only by the true spirit and scope of the appended claims.

We claim: Y

1. In a mirror landing system for landing an aircraft on the deck of aship along a predetermined glide path, wherein said landing systemincludes a mirror and a source of light creating a virtual image in andretiected from said mirror, the center of said mirror and said source oflight being fixed with respect to said deck; the combination therewithof automatic means for tilting said mirror about its center through anangle ,u equal to .wherein o is the angle between a rst line determinedby the center of said mirror and an aircraft located on the glide pathand the projection of the first line on a horizontal plane passingthrough the center of said mirror,

'is the pitch angle of said ship,'and 0 is the angle between a secondline connecting said source of light and the center of said mirror andthe projection of said second line in a plane parallel to said deckpassing through saidrsource of light, said automatic means includingtracking means having a sensing device the center of which is fixedrelative to said deck for determining the actual angle a between a thirdline connecting the center of said sensing device and said aircraft andthe projection of said third line in a plane parallel to said deck andpassing through the center of said sensing device.

2. The combination defined in claim 1, wherein said sensing device isresponse to infra-red radiation.

3. The combination defined in claim l, wherein Where Ah is thedierential vertical distance between the actual position of the centerof said mirror and the position the center of said mirror would have ifthe pitch, roll and heave of said ship were all zero, b is the fixeddistance between the centers of said sensing element and said mirror, yis the fixed angle between b and the projection of b in a plane parallelto said deck and passing through the center of said sensing device, andqt' is the angle between said predetermined glide path and theprojection of said glide path in a horizontal plane; and wherein saidautomatic means further includes computer means coupled to said trackingmeans for producing an output manifesting the angle n in response to aiirst input manifesting a second input manifesting a third inputmanifesting 0, a fourth input manifesting qt and a fifth inputmanifesting Ah, and means responsive to the output of said computermeans for tilting said mirror vthrough an angle p..

4. The combination defined in claim 3, wherein said automatic meansfurther includes a height change computer for determining Ah.

5. The combination defined in claim 4, wherein said height changecomputer includes a vertical stabilized vertical accelerometer, andintegration means for determining the double integral of the output ofsaid accelerometer, whereby the output from said integrating meansmanifests Ah.

References Cited in the file of this patent UNITED STATES PATENTS1,238,503 Fiske et al. Aug. 28, 1917 1,558,567 Schein Oct. 27, 19252,784,925 Goodhart Mar. 12, 1957 OTHER REFERENCES Selsyn Drive CatalogueBulletin 7 22, GEA, May 1929,

